1. Field of the Invention
The present invention relates to fluid reaction surfaces, and more specifically to an air cooled turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, either an aero engine or an Industrial Gas Turbine (IGT), a compressor supplies compressed air into a combustor to be mixed with a fuel and burned to produce a hot gas flow through the turbine of the engine. The hot gas flow passes through a multiple stage turbine having a plurality of stationary vane or nozzle stages with an equal number of rotating blade stages arranged in an alternating manner. The turbine progressively reduced the hot gas flow temperature by removing mechanical energy from the flow.
One method of increasing the efficiency of the engine is to provide for a higher entry temperature into the turbine. However, the material properties of the first stage vane and blades have temperature limits for use. In order to increase the turbine inlet temperature, internal cooling of the vanes and blades have been used which allow for higher temperatures, and therefore increased efficiency.
The compressed air used to pass through the internal parts of these vanes and blades is usually bled off from the compressor, which reduces the amount of compressed air in which the engine has performed work thereon that is not used in the combustor. This bleed off air from the compressor also reduces the efficiency of the engine. It is thereof an object of designers of air cooled turbine vanes and blades to provide increased cooling of these turbine members while making use of minimum amounts of compressed air in order to improve the engine efficiency.
In order to cool a turbine airfoil in the prior art, cooling air is passed through an internal cooling circuit. An example of this is disclosed in U.S. Pat. No. 6,544,001 B2 issued to Dailey on Apr. 8, 2003 entitled GAS TURBINE ENGINE SYSTEM in which an airfoil includes a single hollow portion that forms an internal cooling air passage. A plurality of cooling holes discharges cooling air from the hollow portion to the external surface of the airfoil. One major problem with this type of cooling circuit is that a single pass through the airfoil is used, and therefore heat transfer to the cooling air is minimal. Another problem is that the pressure side cooling holes require higher pressure to discharge to the external surface than does the cooling holes of the suction side. In this patent, the air pressure in the hollow portion must be high enough to discharge enough cooling air onto the pressure side, resulting in excess amount of cooling air to be discharged through the suction side cooling holes. Cooling air is wasted, resulting in lower engine efficiency.
To provide for a longer cooling air flow path within an airfoil, the prior art made use of a serpentine cooling flow circuit. U.S. Pat. No. 7,014,424 B2 issued to Cunha et al on Mar. 21, 2006 entitled TURBINE ELEMENT discloses a turbine airfoil with three separated cooling circuit within the airfoil. A first cooling circuit is located in the leading edge portion and discharges cooling air from a channel into a showerhead arrangement to cool the leading edge. A five-pass serpentine circuit is located at the mid-region of the airfoil. Cooling air is supplies in the first leg of the five-pass serpentine circuit and flows upward from root to tip in order that the fifth leg also flows upward in the airfoil to discharge into the airfoil tip. A third separate cooling circuit is located in the trailing edge region. One problem with the cooling circuit of the Cunha et al patent is that the five-pass serpentine circuit is used to cool both the pressure side wall and the suction side wall. In order to provide adequate cooling for the hotter pressure side wall, higher pressure is required and more cooling than is required is used on the suction side wall. Also, the second and fourth legs of the serpentine circuit supply cooling air to cooling holes on both sides of the airfoil. This also results in over-pressure for the suction side and a waste of cooling air discharged onto the suction side. Lower engine efficiency is a result.
To improve on the cooling circuit like the one shown in the Cunha et al patent, some prior art make use of two serpentine circuits at the mid section of the airfoil. U.S. Pat. No. 5,813,835 issued to Corsmeier et al on Sep. 29, 1998 entitled AIR-COOLED TURBINE BLADE shows a prior art cooling circuit (FIG. 3a in this patent) that has a three-pass serpentine circuit on the pressure side and another three-pass circuit on the suction side opposite to the circuit on the pressure side. A divider wall (212 in this patent) separates the two serpentine circuits. The improvement in this cooling circuit is that the suction side serpentine cooling circuit can operate at a lower pressure than the pressure side serpentine circuit, thus requiring less cooling flow to be wasted and therefore improving the engine efficiency. On problem with this cooling circuit is that the both serpentine cooling circuit flow through the passages from the leading edge toward the trailing edge. The Corsmeier et al patent is an improvement to this circuit, in which a third middle cooling circuit is added and positioned between the pressure and suction side cooling circuits. The reason for this is that the divider wall of the prior art cooling circuit tends to be overcooled by the flow of cooling air passing through the serpentine flow passages that surround the divider wall. If the middle portion of the airfoil is overcooled, then thermal gradients occur within the airfoil and produce undesired stress levels. in the Corsmeier et al patent, the flow path of both mid-airfoil serpentine circuit is from trailing edge toward the leading edge. Cooling air is wasted in the pressure side serpentine circuit because of this. The highest pressure acting on the pressure side is near the forward most leg of the pressure side serpentine circuit. The cooling air must flow through the first and second legs of the serpentine circuit in order to reach the third leg and be discharged out through the cooling hole to cool the hottest section of the pressure side wall. Thus, an overpressure is required to supply an adequate amount of cooling air at the necessary pressure for this cooling hole.
The U.S. Pat. No. 6,705,836 B2 issued to Bourriaud et al on Mar. 16, 2004 entitled GAS TURBINE BLADE COOLING CIRCUITS discloses a turbine blade cooling circuit having five independent cooling circuits within the blade (labeled A through E in this patent). Circuit A is a three-pass serpentine circuit on the pressure side and flows in a direction from back to front of the airfoil. Circuit B is a three-pass circuit with two first legs and flows in a back to front direction, opposite to the pressure side serpentine circuit. Circuit C is a leading edge circuit, Circuit D is a trailing portion circuit, and Circuit E cools the trailed edge. The cooling circuits of the Bourriaud et al patent are a near-wall cooling design. A central cavity (6 in this patent) is positioned between the pressure and suction side cooling circuits, and supplies cooling air to the leading edge cavity (* in this patent) of the leading edge cooling circuit C. because of the central cavity, the inner walls of the airfoil are also overcooled as in the above divider wall described in the Corsmeier et al patent. Therefore, thermal gradients occur within the blade and result in undesirable stresses.
It is therefore an object of the present invention to provide for an internal cooling circuit for a turbine airfoil that provides adequate cooling, minimal cooling flow, and provides for a more even temperature distribution throughout the airfoil to reduce stress levels from a thermal gradient.